1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically a turbine rotor blade with platform cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Turbine rotor blades have a platform that forms a flow path for a hot gas stream and thus must also be cooled in order to prevent hot spots that lead to erosion or other damage to the blade. FIG. 1 shows one prior art blade 10 platform cooling design in which the platform includes a pressure side (P/S) and a suction side (S/S) with each having a series of straight cooling channels formed within the platform to provide cooling. The P/S platform includes three straight large channels 12 that discharge on the aft side of the platform. The S/S platform includes one large straight channel 12 that feeds into three smaller straight channels 13 that also discharge out the aft side of the platform. For each of these straight channels 12, a row of cooling air inlet holes are arranged on the forward end of the platform and connect to a dead rim cavity located below the platform for a supply of cooling air. FIG. 2 shows a cross section view through the line A-A in FIG. 1 with the airfoil extending from the platform 11 and three of the straight cooling channels 12 on the P/S platform.
In the prior art blade platform cooling design of FIG. 1, for an airfoil with a low cooling flow design, especially a low platform cooling flow design, better cooling of the platform would require adding more of the straight cooling channels by making them closer together. Using more of these straight cooling channels would require more cooling flow in the platform. Without better cooling of the platform, hot spots will occur in-between the straight cooling channels and produce uneven cooling for the platform surface.